This invention relates to a method and apparatus for correcting distortion of gas turbine engine compressor blades and vanes (hereafter referred to as blades) primarily in alloys with low temperature creep characteristics, such as titanium alloys.
A widely used method for fabricating blades is by precision forging the airfoil and dovetail platform to net shape, heat treating and finally machining the dovetail. Nearly all blades passing through this process develop unacceptable distortion of the airfoil and its relationship to the platform. The predominant proir art method for correcting this distortion is by manually shaping the blade using special hand and gripping tools to strain into tolerance any particular area of the blade structure which is out of tolerance. Because of the complexity of the blade geometry and the lack of accuracy in this manual procedure correcting one area of the blade often results in adverse effect to other areas requiring additional shaping of the blade. This method of shaping blades is time-consuming and requires a highly skilled technician. A further disadvantage of this manual shaping process is that the mechanical deformation of the blades causing residual stresses to be built up in the blade which reduces fatigue strength and makes them highly unstable. Peening is then necessary to alleviate these detrimental stresses.
Another disadvantage of the prior art techniques for manually shaping gas turbine engine blade structures is that such techniques have proved considerably more difficult in correcting distortions in extermely resilient high temperature materials such as titanium alloys which are often used in blade construction. Because of their high resiliency titanium alloy blades tend to return to their original shape when subjected to the application of mechanical deformation. Consequently, manual force often fails to correct distortions in such blades resulting in a high rejection rate for titanium blades and a consequent increase in the production costs of such blades.
It is thus obvious that the conventional manual process for shaping gas turbine engine blades is a quite costly, time-consuming, and unsatisfactory technique for reshaping highly resilient high temperature alloys.
Applicant has found that these disadvantages may be overcome by the use of creep forming techniques. Prior art processes for creep forming generally comprise pressuring a specimen between two opposed dies, elevating the temperature of the specimen to the creep range of the material of which it is constructed and maintaining the die pressure until sufficient creep has transpired as illustrated in FIGS. 1 and 2. Such prior art process for creep forming has been found to be unsatisfactory for correcting distortion in gas turbine engine blades. Several problems are encountered when trying to reshape gas turbine engine blades when using opposed dies. One major problem is that due to blade surface irregularities the pressure load on the die is extremely non-uniform resulting in extremely concentrated pressure loads at various points on the surface of the die which correspond to unusually thick proportions of the blade, as best seen in FIG. 2 where high pressure points are shown generally at 3 and low pressure points are shown generally at 5. These concentrated pressures often result in deformation of the dies.
A further problem with such prior art opposed die creep forming techniques is that some blade configurations do not lend themselves to a two-die system. The geometry of these blades is such that when the blades are compressed between opposed dies significantly greater loading is created in some areas of the die than in others. This uneven loading prevents removal of distortion without creating excessive loading on the dies.
It has also been proposed to use single-die techniques for creep forming such as by use of an autoclave. In such prior art methods, gas or hydraulic pressure is used to apply load to a workpiece abutting a single die cavity. However, because of the method of applying pressure and the relatively expensive apparatus required therefore, such prior art systems have been unable to economically achieve the high production rates required when creep forming small parts such as gas turbine engine blades.